Rocket powered launch vehicles provide the thrust for launching and delivering a payload from the earth into earth orbit. Payloads may include complex assemblies such as satellites. Launch vehicles typically include one or more rocket engines configured to activate at different times or stages as the launch vehicle travels from the earth into orbit. A rocket engine in a launch vehicle may incorporate a solid or liquid propellant contained in a propellant tank, a combustion chamber, and a nozzle for directing the combusted propellant to accelerate the rocket in flight. Liquid propellant rockets are known as bi-propellant rocket systems because the liquid fuel also requires a liquid oxidizer to be stored in a separate tank. The liquid fuel and oxidizer are mixed with one another in desired ratios in order to most efficiently combust the propellant to generate an expanded gas that is accelerated through the nozzle for propelling the launch vehicle.
In order to maximize the efficiency of the propellant burn, a control loop is provided that monitors the amounts of the liquid oxidizer and the liquid fuel remaining in the respective tanks. Ideally, the liquid fuel and oxidizer are metered through valves leading into the combustion chamber at exact desired ratios in order to achieve the most efficient burn of the propellant and ensuring that all the propellant is consumed by the end of flight. Excess of one of the propellant commodities at the end of flight is known as propellant outage (dead weight) and degrades vehicle performance. The purpose of a propellant monitoring system is to minimize outage and provide sufficient control authority to counter vehicle anomalies that could jeopardize mission success. Failure to properly control the mixture ratio of the liquid fuel and oxidizer can result in premature engine shutdown that may ultimately sacrifice the ability of the launch vehicle to place the payload into the targeted orbit.
One general configuration for monitoring the levels of the fuel and oxidizer includes the use of pressure transducers that measure the pressure differential, referred to as the “delta pressure”, within the tanks. The delta pressure is due to the liquid hydrostatic pressure within the propellant tank. By knowing the vehicle acceleration and liquid density, it is possible to back calculate the remaining propellant quantities from the delta pressures. Thus, these pressure readings are provided as inputs to a controller that calculates the most optimum fuel and oxidizer mixture ratios, and then provides output signals to control the operation of various mixing valves and other hardware components that control the amount of liquid fuel and oxidizer delivered to the combustion chamber of the engine.
One reference that discloses a propellant utilization system for a space vehicle capable of controlling mixture ratios for thrust sources of space vehicles includes the U.S. Pat. No. 6,631,314. This reference more particularly discloses a propellant utilization system for a space vehicle that may have first and second thrust sources, for example, a booster stage to launch and deliver a payload from a distance from the earth, and an upper stage that is activated to deliver the payload the remainder of the distance into a desired orbit. This system utilizes a set of algorithms to generate mixture ratios for each thrust source as each thrust source becomes active. The propellant utilization system includes a processing system including sequential logic, propellant logic, and mixture ratio logic. Sequential logic determines when a thrust source is active and provides flight parameters for the active thrust source to the propellant logic and the mixture ratio logic. The propellant logic processes information from propellant sources connected to the active thrust source, using the flight parameters for that thrust source to determine an amount of remaining propellant in each source. The mixture ratio logic generates a mixture ratio for the active source, using the flight parameters for that thrust source and information on the remaining amount of propellant in each source connected to the active thrust source. This U.S. Pat. No. 6,631,314 is hereby incorporated by reference in its entirety.
The liquid fuel and oxidizer are contained in tanks with some amount of ullage space above the liquid surface that contains a pressurized gas. The ullage increases as the propellant is consumed. The ullage pressure is controlled during flight so that adequate pressure can be applied to the liquid surface to satisfy the engine feed system requirements at the engine pump. In order to measure the amounts of the oxidizer and liquid propellant within the respective tanks, one method is to use a delta pressure transducer that receives two pressure sensing inputs, one from the liquid or head side (tank bottom) of the tanks and the other from the ullage side (tank top) of the tanks. The differential pressure sensed by pressure transducer at these two points in the tanks enables determination of the remaining amounts of liquid fuel and oxidizer within the respective tanks.
Presently, one method of sensing the pressure on the liquid side of the tanks is by the use of a small diameter tube that has one end immersed in the liquid, referred to as the head pressure sensing line. The head pressure sensing line utilizes purge gas to prevent entry of liquid within the sense line. Since the terminus of the sensing tube is immersed within the liquid, the column of liquid above the open end of the sensing tube will exert a certain amount of pressure which is transferred through the gas in the sensing line to the pressure transducer. Because of the extreme conditions under which the space vehicle operates, liquid can be forced into the sense lines by fluid pressure fluctuations generated primarily by the vibrations of the vehicle structure, acceleration transients, ullage pressurization transients, and other disturbances. Some of these disturbances can be unpredictable. More specifically, when pressure fluctuations in the liquid are of sufficient amplitude and frequency, liquid is ingested within the sense lines, but the small diameter of the sense lines make it more difficult for the liquid to be removed from the lines despite the presence of the expelling force from the purge gas. When liquid is ingested in a head sense line, the liquid prevents the purge gas from flowing out of the line at a normal rate. The purge gas continues to flow into the sense line from its source thereby increasing the pressure within the head sense line. The increased pressure in the head sense line results in an incorrect measurement of pressure in the line that is higher than the actual liquid head pressure within the tank. Repeated pressure fluctuations result in repeat slugs of liquid ingestion into the head sense line that result in a mean pressure offset (bias) due to the build up of purge gases in the sense line. This bias cannot be removed by electronic filtering or other signal processing techniques. An intermittent or continued incorrect pressure reading on either of the tanks due to liquid ingestion in the head sense lines corrupts the propellant utilization system's ability to accurately determine the actual fuel/oxidizer levels, therefore resulting in erroneous mixture ratio adjustments. The erroneous mixture ratio adjustments can result in a propellant imbalance at engine shutdown causing a loss or reduction in vehicle payload delivery performance.
Therefore, there is a need for an improved device and system that prevents liquid ingestion into a pneumatic pressure sensing line. Preferably, the solution is one that has features of simplicity, will work well with cryogenic rocket propellants such as liquid oxygen, can be retrofit into existing systems without major redesign, has no moving parts, low weight, and can be tuned to filter-out specific pressure disturbances.